The turbines used in modern gas turbine engines are required to operate at extremely high temperatures. In order for the aerofoil blades or vanes present in those turbines to withstand such high temperatures, it is necessary to cool them. This is typically achieved by providing the blades or vanes with internal passages, through which a cooling fluid, usually air, can be passed.
In order to maximise the efficiency of heat transfer from a blade or vane to the cooling fluid, a single passage may pass through the blade or vane several times. This will inevitably mean that the passages have bends around which the cooling fluid must flow. Unfortunately, as the cooling fluid flows round the bends, it experiences a drop in pressure, which can be particularly large where a bend subtends a large angle (eg 180°). Such pressure drops can be problematic if, for example, the cooling fluid is subsequently required for film cooling of an external surface of the blade or vane. In addition, extra pressure loss may necessitate an increase in coolant pressure, which can in turn increase leakage in the system and directly affect engine cycle efficiency. Film cooling involves the fluid being exhausted through a plurality of small holes connecting the internal cooling passages with the blade/vane exterior. Any loss in pressure when traversing the bends in the passages will reduce the amount of fluid that can be exhausted to the blade/vane exterior, so reducing the overall film cooling.
Several methods of reducing the loss in pressure caused by bends in the cooling passages have been suggested. One example is to provide turning vanes in the bends. Although these reduce the overall pressure loss, they increase the weight of the blade or vane and its manufacturing complexity. Another possibility is to vary the shape of a wall member that divides the two passage portions on either side of a bend. U.S. Pat. No. 5,073,086 describes an aerofoil blade having pressure and suction flanks, in which the flanks are interconnected internally of the aerofoil portion by a generally longitudinally extending wall member to partially define first and second cooling fluid passage portions. The first and second passage portions are interconnected in series fluid flow relationship by a bend passage portion, and the wall member is locally thickened in the region of the bend passage portion to provide a localised progressive series narrowing and opening of the upstream end of the second passage portion in the general direction of cooling fluid flow.
However, there is a continuing need to minimise the loss in pressure experienced when the coolant flows round a bend in the cooling passage.